This disclosure generally relates to a gas turbine engine system.
In an aircraft gas turbine engine, such as a turbofan engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. The hot combustion gases flow downstream through turbine stages and extract energy from the gas. In a two spool gas turbine engine, a high pressure turbine powers the high pressure compressor, while a low pressure turbine powers a fan section disposed upstream of the compressor and a low pressure compressor.
Combustion gases are discharged from the turbofan engine through a core exhaust nozzle and fan air is discharged through an annular fan exhaust nozzle defined at least partially by a nacelle surrounding the core engine. A majority of propulsion thrust is provided by the pressurized fan air which is discharged through the fan exhaust nozzle, while the remaining thrust is provided from the combustion gases discharged through the core exhaust nozzle.
It is known in the field of aircraft gas turbine engines that the performance of the engine varies during diverse flight conditions experienced by the aircraft. An inlet lip section located at the foremost end of the nacelle is typically designed to enable operation of the engine and prevent the separation of airflow from the inlet lip section of the nacelle during diverse flight conditions. For example, the inlet lip section requires a “thick” inlet lip section design to support operation of the engine during specific flight conditions, such as cross-wind conditions, take-off and the like.